The assignee of the present invention manufactures and deploys spacecraft for communications and broadcast services. To meet market demands for advanced services from such spacecraft, spacecraft payloads of increased size and improved pointing performance are required. For example, there is a demand for increased aperture antenna reflectors, having diameters of three meters or greater, and larger, more complex, radio frequency (RF) feed arrays. A payload element, such as, for example, an antenna reflector, and/or an RF feed array, often require one or more positioning mechanisms configured to provide for (i) initial deployment of the payload element from a stowed launch position to an (on-orbit) operating position, and/or (ii) on-orbit steering of the reflector or feed to provide precise pointing.
Such positioning of spacecraft payload elements has conventionally been performed, by, for example, linear or rotary actuators as illustrated in FIG. 1A and FIG. 1B, respectively. Referring to FIG. 1A, spacecraft main body structure 100 is coupled to reflector 120 by way of panel 110. Panel 110 may have a deployment hinge (not shown) proximate to spacecraft 100 to facilitate an initial deployment from a stowed (launch) configuration to a deployed (on-orbit) configuration. Operation of linear actuators 132 results in rotation of reflector 120 about each of two mutually orthogonal axes 136 and 138 defined by pivots 134. Referring now to FIG. 1B, rotation of reflector 120 in each of two mutually orthogonal axes 136 and 138 may, alternatively, be accomplished by way of two rotary actuators 140.
Positioning of spacecraft antenna reflectors and other appendages using the techniques described above, and variants thereof, has been used successfully. For spacecraft requiring larger reflectors and/or more stringent pointing requirements, however, such techniques are problematic. For example, in the rotary actuator implementation illustrated in FIG. 1B, rotary actuators 140, located near the spacecraft main body structure 100, are at a large distance from the center of gravity (c.g.) of reflector 120. This results in a low deployed natural frequency, and, consequently pointing performance degradation, because the moment of inertia of reflector 120 is large, particularly about axis of rotation 138.
Locating a positioning mechanism behind reflector 120 (nearer the c.g.) is sometimes possible, as illustrated in FIG. 1A. However, this increases the inertia about the deployment hinge and puts the actuators in a severe thermal environment. In addition, for certain types of unfurlable mesh reflectors, antenna backup structure 121, shown schematically in FIG. 1A, may be absent, or incompatible with an interface to linear actuators 132 and pivots 134.
Moreover, locating the actuators behind or at an edge of reflector 120, results in rotation of reflector 120 about a point substantially distant from an RF focal point of reflector 120, resulting in defocusing of payload beam and consequent performance degradation.
In light of the foregoing problems, improved payload positioning mechanisms are desirable.